Braze repair of a gas turbine engine stationary shroud

ABSTRACT

An undersize repair region of a gas turbine engine stationary shroud is repaired with a sufficient mass of a repair material. The repair material includes a first fraction of a first powder of a first alloy component, and a second fraction of a second powder of a second alloy component. The first alloy component and the second alloy component have different solidus temperatures. The repair material is placed in the repair region. The repair material and the repair region are heated to melt the repair material but not the repair region, and thereafter the repair material and the repair region are cooled to solidify the repair material.

This application is a continuation-in-part of application Ser. No.09/322,008, filed May 28, 1999, now U.S. Pat. No. 6,283,356.

FIELD OF THE INVENTION

This invention relates to gas turbine engines and, more particularly, tothe repair of stationary shrouds found in gas turbine engines.

BACKGROUND OF THE INVENTION

In an aircraft gas turbine (jet) engine, air is drawn into the front ofthe engine, compressed by a shaft-mounted compressor, and mixed withfuel. The mixture is burned, and the hot combustion gases are passedthrough a turbine mounted on the same shaft. The flow of combustion gasturns the turbine by impingement against an airfoil section of theturbine blades and vanes, which turns the shaft and provides power tothe compressor. The hot exhaust gases flow from the back of the engine,driving it and the aircraft forwardly.

The turbine blades are mounted on a turbine disk, which rotates on ashaft inside a generally cylindrical tunnel defined by a hollowstationary shroud structure. The stationary shroud structure is formedof a series of stationary shrouds that extend around the circumferenceof the tunnel in an end-to-end fashion. The stationary shroud structurehas such a segmented arrangement to accommodate the thermal expansionexperienced during each engine cycle as the stationary shroud structureis cycled between room temperature and a maximum service temperature ofover 2000° F. Each of the stationary shrouds has an internal gas pathsurface that is a segment of a cylinder, and a support structure thatbacks the gas path surface and provides for attachment to the adjacentstructure.

During service, the support structure of the shrouds may be damaged byfatigue, erosion, and other mechanisms. One form of the damage is thewearing away of material from the shrouds, at locations such as the endfaces, the forward and aft edges, and elsewhere. As material is wornaway and during multiple repair cycles when material is removed bymachining operations, the shroud gradually becomes undersize in at leastone dimension of the support structure. When the shroud has become toosmall in at least one dimension of the support structure to continue tobe functional, it is discarded.

There is a need for an improved approach to responding to such damage togas turbine engine shrouds. The shrouds are made of expensivenickel-base or cobalt-base superalloys, and the discarding of a shroudrepresents a substantial cost. The present invention fulfills this need,and further provides related advantages.

BRIEF SUMMARY OF THE INVENTION

The present invention provides a method of repairing a gas turbineengine stationary shroud. The repair may be performed on any portions ofthe support structure. It is preferably performed on the end faces whichbutt against the end faces of the neighboring shrouds in service andgradually become undersized. The repaired shroud is fully functional andis serviceable at a small fraction of the cost of a new shroud.

A method of repairing a gas turbine engine stationary shroud comprisesthe steps of providing the gas turbine engine stationary shroud havingan undersize repair region made of a shroud material, wherein the repairregion is not located on a gas flow path surface of the gas turbineengine stationary shroud. The repair region may be, for example, an endface, an edge, or a back surface of the gas turbine engine stationaryshroud. The repair region of the gas turbine engine stationary shroud isrepaired so that the repair region is no longer undersize. The step ofrepairing includes the steps of providing a sufficient mass of a repairmaterial comprising a first fraction of a first powder of a first alloycomponent, and a second fraction of a second powder of a second alloycomponent. The first alloy component and the second alloy component havedifferent solidus temperatures. Each of the two powders is preferablyprealloyed, so that its constituents are melted together prior to thetwo powder types being mixed together. The step of repairing furtherincludes placing the repair material in the repair region, heating therepair material and the repair region to a brazing temperaturesufficient to melt the repair material but not the shroud material ofthe repair region, so that the repair material flows over the repairregion, and thereafter cooling the melted repair material and the repairregion to solidify the repair material, the repair material having asolidus temperature less than that of the shroud material.

The shroud material may be a cobalt-base superalloy or a nickel-basesuperalloy, and the repair material is selected accordingly. The firstpowder and the second powder that form the repair material may beprovided as free-flowing powders, or they may first be mixed andsintered together to form a pre-sintered compact. The use of thepre-sintered compact is preferred for standard repair locations, such asfor use at the end faces.

The present approach achieves a fully serviceable repaired shroud,reducing the number of shrouds that are discarded. Other features andadvantages of the present invention will be apparent from the followingmore detailed description of the preferred embodiment, taken inconjunction with the accompanying drawings, which illustrate, by way ofexample, the principles of the invention. The scope of the invention isnot, however, limited to this preferred embodiment.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a perspective view of a turbine blade positioned adjacent to ashroud structure;

FIG. 2 is a perspective view of a single shroud from the supportstructure side; and

FIG. 3 is a block flow diagram of a method for repairing the shroud.

DETAILED DESCRIPTION OF THE INVENTION

FIG. 1 depicts a turbine blade 20 mounted to a periphery 22 of a turbinedisk 24. There are a large number of turbine blades 20 mounted in thisfashion to the turbine disk 24, but only one is illustrated. The turbinedisk 24 rotates on a turbine shaft (not shown) positioned along itscenterline. As the turbine disk 24 rotates, the turbine blade 20 sweepsthrough an annular volume between the turbine disk 24 and a stationaryshroud structure 26, a portion of the circumference of which is shownschematically in FIG. 1 and in more detail in FIG. 2. The shroudstructure 26 in its entirety defines a tunnel 28 in which the turbinedisk 24, turbine shaft, and turbine blades 20 rotate. Hot combustiongases flow from a combustor (not shown) through the annular volume ofthe tunnel 28 between the periphery 22 of the turbine disk 24 and theshroud structure 26, impinging against the turbine blades 20 and causingthe turbine disk 24 and the shaft to turn.

The shroud structure 26 is formed of a number of individual shrouds 30positioned in an end-to-end arrangement around the circumference of thetunnel 28. One of the shrouds 30 is shown in greater detail in FIG. 2.The shroud 30 has a gas flow path surface 32 (the underside of theshroud 30 in the view of FIG. 2) which faces the turbine blade 20.

A support structure 34 forms the back side of the shroud remote from thegas flow path surface 32. The support structure 34 includes oppositelydisposed end faces 36 that abut the end faces of the adjoining shrouds30, and oppositely disposed forward edge 38 and aft edge 40. Additionalstructural features, whose details and functions are not pertinent tothe present invention, include a forward groove 42, an aft groove 44, arace track 46, casting ribs 48, and a back surface 50.

During service, one or more of the features of the support structure 34may become damaged by removal of metal, so that it becomes undersized.Initially, some such damage is acceptable, but eventually the featurebecomes so far below its desired specified service minimum dimensionthat it is no longer functional. In the past, it has been the practiceto discard the entire shroud at this point. The present inventionprovides a repair technique for the support structure 34 of the shroud30 so that it may be removed from the engine, repaired, and thenreturned to service.

FIG. 3 illustrates a preferred method for performing the repair. A gasturbine engine stationary shroud 30 is provided, numeral 60. The shroud30 has a repair region which is undersize. That is, some dimension ofthe shroud 30 is less than a specified service minimum dimension. Thecurrent repair region of most concern, which will be discussed in detailfor the sake of definiteness, is loss of material from the end faces 36that abut the end faces of the neighboring shrouds. As illustrated inFIG. 2, a specified service minimum dimension D, the width of thecasting rib 48, is indicative of the total chord length of the shroud 30between the oppositely disposed end faces 36. If D is too small, theshroud 30 will be too short in the circumferential direction and willnot fit together properly with the adjacent shrouds, allowing turbinegas leakage between the shrouds and a resulting decrease in operatingefficiency.

The shroud 30 and its repair region, in this case the repair region 36,are repaired by a technique involving activated diffusion healing brazerepair, numeral 62. A repair material is provided, numeral 64. Therepair material is a sufficient mass of a first fraction of a firstpowder of a first alloy component, and a second fraction of a secondpowder of a second alloy component, to restore the repair region back toits desired dimension. The first alloy component and the second alloycomponent have different solidus temperatures. The repair material thatis later formed as a melted mixture of the first powder and the secondpowder has a solidus temperature less than that of a shroud materialthat forms the repair region.

The first powder and the second powder are selected according to theshroud material that forms the repair region. The powders selected forcobalt-base shroud materials are different from those selected fornickel-base shroud materials. In a case of particular interest, theshroud material is a cobalt-base alloy known as Mar M509, which has anominal composition, in weight percent, comprising about 23.5 percentchromium, about 10 weight percent nickel, about 7 percent tungsten,about 3.5 percent tantalum, about 0.2 percent titanium, about 0.4percent zirconium, about 0.6 percent carbon, no more than about 2percent iron, balance cobalt and impurities.

For the cobalt-base shroud material, the first alloy componentpreferably comprises a prealloyed composition, in weight percent, offrom about 10 to about 25 percent nickel, from about 15 to about 25percent chromium, from about 5 to about 10 percent silicon, from about 2to about 6 percent tungsten, from about 0.2 to about 0.8 percent carbon,from about 0.4 to about 2.0 percent boron, balance cobalt andimpurities. The second alloy component preferably comprises a prealloyedcomposition, in weight percent of from about 5 to about 15 percentnickel, from about 15 to about 30 percent chromium, about 2.0 percentmaximum silicon, from about 5 to about 10 percent tungsten, from about0.3 to about 0.8 percent carbon, about 1.5 percent maximum manganese,about 3 percent maximum iron, about 0.5 percent maximum zirconium,balance cobalt and impurities. The first fraction is preferably fromabout 25 weight percent to about 50 weight percent, most preferablyabout 35 weight percent. The second fraction is preferably from about 75weight percent to about 50 weight percent, most preferably about 65weight percent.

On the other hand, the shroud material may be a nickel-base superalloysuch as Rene N5, which has a nominal composition, in weight percent, offrom about 6 to about 6.4 percent aluminum, from about 6.75 to about7.25 percent chromium, from about 7 to about 8 percent cobalt, fromabout 0.12 to about 0.18 percent hafnium, from about 1.3 to about 1.7percent molybdenum, from about 2.75 to about 3.25 percent rhenium, fromabout 6.3 to about 6.7 percent tantalum, from about 4.75 to about 5.25percent tungsten, a sum of aluminum plus tantalum about 12.45 percentminimum, balance nickel and impurities. Where the shroud material is anickel-base superalloy such as Rene N5, the first alloy componentpreferably comprises a prealloyed composition, in weight percent, offrom about 10 to about 20 percent cobalt, from about 14 to about 25percent chromium, from about 2 to about 12 percent aluminum, from 0 toabout 0.2 percent yttrium, balance nickel and impurities. The secondalloy component preferably comprises a prealloyed composition, in weightpercent of from about 10 to about 20 percent cobalt, from about 14 toabout 25 percent chromium, from about 2 to about 12 percent aluminum,from about 2 to about 12 percent silicon, balance nickel and impurities.The first fraction is preferably from about 55 to about 80 weightpercent, most preferably about 68.5 weight percent. The second fractionis preferably from about 45 weight percent to about 20 weight percent,most preferably about 31.5 weight percent.

The two types of individually prealloyed powders may be provided in aloose, free-flowing form. They may instead be provided as a pre-sinteredcompact. Both approaches are operable, although the use of thepre-sintered compact is more practical for production operations. Inthis latter approach, the powders are mixed together, pressed with abinder into a desired shape, and sintered by heating to a temperaturewhere the powders sinter but do not both melt. It is not necessary thatthe pre-sintered compact have a high relative density (that is, littleporosity), as it is later melted. The pre-sintered compact is moreeasily handled and positioned than are the free-flowing powders, andthere is less compaction and shrinkage in subsequent melting. Acombination of these approaches may be desired. For example, thepre-sintered compact may be contacted to the end face 36, andfree-flowing powders may be packed into the adjacent portions of thegrooves 42 and 44.

The repair material is placed into the repair region, numeral 66. Therepair material may be the mixture of the free-flowing powders, thepre-sintered compact, or a combination of both approaches. The amount ofrepair material is selected so that, after subsequent melting andmachining, the repair region is restored to its desired dimension.

The repair material and the repair region are heated to a brazingtemperature to melt at least a portion of the repair material but notthe shroud material of the repair region, numeral 68. In the case of theabove-discussed repair materials for the cobalt-base alloys and thenickel-base alloys, the brazing temperature is from about 2190° F. toabout 2335° F., preferably from about 2300° F. to about 2325° F. At thebrazing temperature, the powder having the lower solidus temperaturemelts to accelerate the bonding to the shroud and the densificationprocess, while the powder having the higher solidus temperature remainssolid so that the powder mass generally retains its shape.

After a short time at the brazing temperature, typically on the order ofabout 20 minutes to about 2 hours, preferably about 2 hours, the meltedrepair material and the repair region are cooled below the solidustemperature of the melted repair material to solidify the repairmaterial, numeral 68. The repair material solidifies bonded to theshroud. The result is a shroud 30 in which the repair region is nolonger undersize.

In most cases, the amount of repair material is selected so that therepair region will be oversize after the brazing and cooling steps.Although it would be desirable to make the repair exactly the right sizeafter brazing and cooling, it is typically not possible to control theamount and distribution of the repair metal that precisely. Accordingly,the repair region is made oversize, and then machined, numeral 70, tothe correct size and with the necessary details such as the grooves 42and 44.

The present invention has been reduced to practice. A total of 400-500shrouds 30 made of Mar M509 material were repaired on their end faces torestore their proper chord lengths. Repaired shrouds were tested byoxidation testing at 2050° F. for 23 hours, and by furnace cycle testingbetween room temperature and 2000° F. for 200 cycles. The performance ofthe repaired shrouds was equivalent to that of the original Mar M509substrate material.

Although a particular embodiment of the invention has been described indetail for purposes of illustration, various modifications andenhancements may be made without departing from the spirit and scope ofthe invention. Accordingly, the invention is not to be limited except asby the appended claims.

What is claimed is:
 1. A method of repairing a gas turbine enginestationary shroud, comprising the steps of providing the gas turbineengine stationary shroud having an undersize repair region made of ashroud material, wherein the repair region is not located on a gas flowpath surface of the gas turbine engine stationary shroud; repairing therepair region of the gas turbine engine stationary shroud so that therepair region is no longer undersize, the step of repairing includingthe steps of providing a sufficient mass of a repair material comprisinga first fraction of a first powder of a first alloy component, and asecond fraction of a second powder of a second alloy component, whereinthe first alloy component and the second alloy component have differentsolidus temperatures, placing the repair material into the repairregion, heating the repair material and the repair region to a brazingtemperature sufficient to melt at least a portion of the repair materialbut not the shroud material of the repair region, so that the repairmaterial flows over the repair region, and thereafter cooling the meltedrepair material and the repair region to solidify the repair material,the repair material having a solidus temperature less than that of theshroud material.
 2. The method of claim 1, wherein the step of providingthe gas turbine engine stationary shroud includes the step of providingthe repair region having a dimension less than a specified serviceminimum dimension.
 3. The method of claim 1, wherein the step ofproviding a gas turbine engine stationary shroud includes the step ofproviding a gas turbine engine stationary shroud made of a shroudmaterial comprising a cobalt-base alloy.
 4. The method of claim 3,wherein the cobalt base alloy has a composition, in weight percent,comprising about 23.5 percent chromium, about 10 weight percent nickel,about 7 percent tungsten, about 3.5 percent tantalum, about 0.2 percenttitanium, about 0.4 percent zirconium, about 0.6 percent carbon, no morethan about 2 percent iron, balance cobalt and impurities.
 5. The methodof claim 3, wherein the first alloy component comprises a prealloyedcomposition, in weight percent, of from about 10 to about 25 percentnickel, from about 15 to about 25 percent chromium, from about 5 toabout 10 percent silicon, from about 2 to about 6 percent tungsten, fromabout 0.2 to about 0.8 percent carbon, from about 0.4 to about 2.0percent boron, balance cobalt and impurities, and the second alloycomponent comprises a prealloyed composition, in weight percent of fromabout 5 to about 15 percent nickel, from about 15 to about 30 percentchromium, about 2.0 percent maximum silicon, from about 5 to about 10percent tungsten, from about 0.3 to about 0.8 percent carbon, about 1.5percent maximum manganese, about 3 percent maximum iron, about 0.5percent maximum zirconium, balance cobalt and impurities.
 6. The methodof claim 5, wherein the first fraction is from about 25 weight percentto about 50 weight percent, and the second fraction is from about 75weight percent to about 50 weight percent.
 7. The method of claim 1,wherein the step of providing a gas turbine engine stationary shroudincludes the step of providing a gas turbine engine stationary shroudmade of a shroud material comprising a nickel-base superalloy.
 8. Themethod of claim 7, wherein the first alloy component comprises aprealloyed composition, in weight percent, of from about 10 to about 20percent cobalt, from about 14 to about 25 percent chromium, from about 2to about 12 percent aluminum, from 0 to about 0.2 percent yttrium,balance nickel and impurities, and the second alloy component comprisesa prealloyed composition, in weight percent of from about 10 to about 20percent cobalt, from about 14 to about 25 percent chromium, from about 2to about 12 percent aluminum, from about 2 to about 12 percent silicon,balance nickel and impurities.
 9. The method of claim 8, wherein thefirst fraction is from about 55 to about 80 weight percent, and thesecond fraction is from about 45 weight percent to about 20 weightpercent.
 10. The method of claim 1, wherein the repair region is an endface of the gas turbine engine stationary shroud.
 11. The method ofclaim 1, wherein the repair region is an edge of the gas turbine enginestationary shroud.
 12. The method of claim 1, wherein the repair regionis a back surface of the gas turbine engine stationary shroud.
 13. Themethod of claim 1, wherein the step of providing a sufficient massincludes the step of providing the first powder and the second powder asfree-flowing powders.
 14. The method of claim 1, wherein the step ofproviding a sufficient mass includes the step of providing the firstpowder and the second powder as a pre-sintered compact.
 15. A shroudrepaired according to the method of claim
 1. 16. A method of repairingan article forming a portion of a flow path surface in a gas turbineengine, the article previously exposed to high temperature operation inthe gas turbine engine, the method comprising the steps of: identifyingan undersize dimension of the article, the undersize dimension beingless than a specified minimum service dimension and the undersizedimension being defined at least in part by a non flow path surface ofthe article; providing a repair material comprising a first fraction ofa first powder of an alloy component and a second fraction of a secondpowder of a second alloy component, the first alloy component and thesecond alloy component having different solidus temperatures, whereinthe repair material has a solidus temperature less than that of anarticle material; associating the repair material with the non flow pathsurface of the article; heating the repair material and the article to abrazing temperature sufficient to melt at least a portion of the repairmaterial but not the article; and cooling the melted repair material andthe repair region to solidify the repair material, the repair materialhaving a solidus temperature less than that of the shroud material, thesolidified repair material increasing the undersize dimension.
 17. Themethod of claim 16, wherein the article comprises a stationary componentof a gas turbine engine, and wherein the undersize dimension of thearticle comprises a chord length.
 18. The method of claim 16 wherein thenon flow path surface is an end face of a gas turbine engine stationaryshroud.
 19. The method of claim 16, wherein the first alloy componentcomprises a prealloyed composition, in weight percent, of from about 10to about 20 percent cobalt, from about 14 to about 25 percent chromium,from about 2 to about 12 percent aluminum, from 0 to about 0.2 percentyttrium, balance nickel and impurities, and the second alloy componentcomprises a prealloyed composition, in weight percent of from about 10to about 20 percent cobalt, from about 14 to about 25 percent chromium,from about 2 to about 12 percent aluminum, from about 2 to about 12percent silicon, balance nickel and impurities.
 20. The method of claim16, wherein the first fraction is from about 55 to about 80 weightpercent, and the second fraction is from about 45 weight percent toabout 20 weight percent.